Aircraft electrically-assisted propulsion control system

ABSTRACT

This invention concerns an aircraft propulsion system in which an engine has an engine core comprising a compressor, a combustor and a turbine driven by a flow of combustion products of the combustor. At least one propulsive fan generates a mass flow of air to propel the aircraft. An electrical energy store is provided on board the aircraft. At least one electric motor is arranged to drive the propulsive fan and the engine core compressor. A controller controls the at least one electric motor to mitigate the creation of a contrail caused by the engine combustion products by altering the ratio of the mass flow of air by the propulsive fan to the flow of combustion products of the combustor. The at least one electric motor is controlled so as to selectively drive both the propulsive fan and engine core compressor.

CROSS-REFERENCE TO RELATED APPLICATION

This specification is based upon and claims the benefit of priority fromUK Patent Application Number 1622012.1 filed on 22 Dec. 2016, the entirecontents of which are incorporated herein by reference.

BACKGROUND OF THE INVENTION 1. Field of the Disclosure

The present disclosure concerns aircraft propulsion systems in whichelectrical energy is used to supplement the operation of a combustionengine. The invention particularly concerns methods for aiding in thecontrol of the negative climate impact resulting from operation of acombustion engine, for example via vapour trail formation.

2. Description of the Related Art

Vapour trails, also known as condensation trails or contrails, areline-shaped ice clouds that appear behind aircraft under certaincircumstances. The formation of a contrail depends on a number offactors, including: ambient temperature, humidity and pressure; theefficiency of the aircraft's engines; and the properties of the fuelburned in the engines.

During the day, contrails reflect a proportion of incoming sunlightaway, leading to a climate cooling effect which is at least partiallyoffset against the climate warming impact associated with the absorptionby contrails of heat radiating upward from the planet's surface. Duringthe night, however, the cooling effect is not operative. For thisreason, a contrail's climate warming impact (per unit time of contrailexistence) is typically greater during the night than during the day. Itis estimated that, globally, night-time flying accounts for some 40% oftotal flying miles, but for some 60% of aviation's totalcontrail-related climate warming impact. The climate warming impact of acontrail is also influenced by its temperature such that a contrailforming in colder air will exert more of a climate warming effect thanan otherwise equivalent contrail which forms in slightly warmer air. Acontrail, once formed, will typically dissipate within a minute or so,unless the ambient air is supersaturated with respect to ice, in whichcase the contrail may persist. A persistent contrail may grow over timeto resemble natural cirrus cloud, both in size and optical properties,and is then referred to as “contrail-cirrus”. Line-shaped contrails andcontrail-cirrus are collectively referred to as “aviation-inducedcloudiness” (AIC). Contrail-cirrus is thought to cause a majority of thenegative climate impact of AIC due to it being spatially larger andlonger-lived relative to a non-persistent line-shaped contrail.

Depending on the metric employed, the climate-warming impact of aviationinduced cloudiness may be of a similar magnitude to that of the carbondioxide (CO2) emitted by aircraft, and may therefore represent asignificant element of aviation's total climate impact. The partial orcomplete suppression of contrail formation, and particularly the partialor complete suppression of persistent contrails, may represent anopportunity for a significant reduction in the overall climate warmingimpact of aviation. Here we use the term “partial suppression ofcontrails” to mean a reduction in the optical depth of formed contrails.

However, a number of potential techniques for reduction of contrailformation or modification of contrail properties by a gas turbine enginerequire the use of bespoke equipment and/or materials that areadditional to those required for conventional engine operation. Anyweight and/or energy penalties incurred in order to achieve contrailsuppression or modification of contrail properties require carefulscrutiny to determine whether such penalties outweigh the possiblecontrail reduction benefits on climate impact.

Conventional propulsion systems for civil aircraft typically compriseone or more turbofan engine placed under the wings of the aircraft.However, some studies have indicated that so-called distributedpropulsion, which involves having numerous smaller propulsion unitspreferentially arranged around an aircraft, may provide some significantbenefits in terms of noise reduction and fuel efficiency when comparedwith current state of the art propulsive arrangements.

The distributed propulsion units are typically electrically driven withthe electrical energy being supplied from a power source mounted ontothe airframe. The power source may be a battery but this is usuallysupplemented by an on-board generator powered using a reciprocatingengine or, more usually, a gas turbine. It has been recognised that thisarrangement can have a greater propensity for contrail formation than aconventional turbofan engine since the exhaust from the reciprocating orgas turbine engine has a relatively high water-vapour partial pressurerelative to the temperature of the exhaust i.e. the “contrail factor”,which is used herein to refer to the gradient of a line representing themixing of engine exhaust air with ambient air, when plotted on a chartusing water-vapour partial pressure as the y-axis and temperature as thex-axis, is relatively high. A reduction in an engine's contrail factorreduces the range of ambient conditions under which the engine can forma contrail. Alternatively, at a particular ambient condition(characterised by pressure, temperature and humidity), a reduction incontrail factor may allow a transition from formation of a contrail tonon-formation of a contrail.

It is an object of the invention to manage the formation of contrails byaircraft engines in a manner that reduces negative impact on theclimate. It may be considered an additional or alternative object toprovide a practical system for implementation of contrail control for anaircraft.

BRIEF SUMMARY OF THE INVENTION

According to a first aspect of the invention there is provided anaircraft propulsion system comprising: an engine having an engine corecomprising a compressor, a combustor and a turbine driven by a flow ofcombustion products of the combustor; at least one propulsive fan forgenerating a mass flow of air to propel the aircraft; an electricalenergy store on board the aircraft; at least one electric motor arrangedto drive the propulsive fan and the engine core compressor; and acontroller arranged for control of the at least one electric motor tomitigate the creation of a contrail caused by the combustion products byaltering the ratio of the mass flow of air by the propulsive fan to theflow of combustion products of the combustor, wherein control of the atleast one electric motor comprises selectively driving both thepropulsive fan and engine core compressor.

The controller may monitor conditions indicative of contrail formation.The controller may monitor conditions indicative of the formation ofcontrails having a negative climate impact.

The controller may reduce the flow of combustion products produced bythe engine core to mitigate or suppress the creation of a contrail.

The controller may increase the power/torque supplied to the propulsivefan by the electric motor during contrail mitigation. The controller mayincrease/control the rotational velocity of the propulsive fan via theelectric motor during contrail mitigation.

The controller may control operation of the electric motor so as to meeta thrust demand for the aircraft. Where the flow rate of combustionproducts is reduced by lowering the rate of fuel burn in the engine corethis may reduce torque and/or thrust generated by the engine core. Thereduced torque and/or thrust may be mitigated by increasing the torqueapplied to one or more propulsive fan by the electric motor.

The at least one electric motor may selectively assist the engine corecompressor, e.g. supplementing the torque applied to the compressor dueto the engine core combustion process.

The controller may control the operation of the at least one electricmotor according to one or more operational variable for the engine core,e.g. the engine core compressor.

The controller may monitor operation of the engine core compressor. Thecontroller may monitor an absolute or relative operational variable forthe engine core compressor, such as any or any combination of rotationalspeed, torque, flow rate and/or pressure rise. Additionally oralternatively, the controller may monitor temperature and/or fuel flowrate for the combustor.

The controller may monitor the relative operation between the propulsivefan and the engine core compressor, e.g. according to relative values ofthe operational variable.

The controller may operate the electric motor to selectively drive theengine core compressor according to one or more predetermined thresholdvalue of the operational variable. The threshold may comprise a safetythreshold, such as a compressor surge margin.

The at least one electric motor may be controlled according to a controlhierarchy wherein priority is given to driving the one or morepropulsive fan or the engine core compressor for contrail mitigation,e.g. provided the threshold is not met or exceeded.

The at least one electric motor may comprise a first electric motorarranged to drive the propulsive fan and a second electric motorarranged to drive the engine core compressor. The engine core compressormay be assisted by the second electric motor selectively, e.g. onlyduring operation of the first electric motor or only during a portion ofthe operational range of the first electric motor. Alternatively, asingle electric motor may be provided, with a suitable selectivetransmission arrangement for driving either or both of the propulsivefan and engine core compressor.

The propulsive fan may be driven at least in part by a low pressureturbine, e.g. via a low pressure shaft. The propulsive fan may becoaxial with the engine core.

A gearing may be provided for altering the speed of the propulsive fanrelative to the speed of an intermediate pressure turbine. The systemmay comprise a geared turbofan engine.

The engine may comprise an intermediate compressor (IPC), e.g. havingone or more row of compressor blades, for compressing air to the enginecore compressor. The propulsive fan may be drivingly connected to theIPC via a gearing.

A gearing may be provided between the at least one electric motor andthe propulsive fan and/or engine core compressor.

A propulsive fan rotor may have one or more row of low pressurecompressor blades for compressing air to the engine core compressor. Thepropulsive fan may comprise a booster. The compressor/booster may rotatein unison, e.g. at a corresponding rotational speed, with the propulsivefan.

The engine may comprise a two-shaft gas turbine engine. The engine maycomprise a boosted turbofan engine.

The system may comprise one or more ambient condition sensor, e.g. forsensing any or any combination of ambient temperature, pressure and/ormoisture/ice content.

The system may comprise one or more electrical generator for chargingthe electrical energy store. The generator may be driven by the engine.The first and/or second motor may operate as a generator, e.g. when notbeing used to drive the respective propulsive fan or engine corecompressor. Accordingly, the term ‘electric motor’ as used herein may beconsidered to be a reference to an electric machine, including generatorfunctionality, where applicable.

There may be one or more further propulsive units which are distributedaround the fuselage and/or wings of the aircraft. The further propulsiveunits may be electrically driven fans connected to the electrical energystore via an electrical network, wherein upon sensing conditionsindicative of contrail formation power supply to the further propulsiveunits from the electrical energy store may be initiated or adjusted,e.g. as either a power increase or decrease.

According to a second aspect of the invention, there is provided amethod of operating an aircraft propulsion system having an engine withan engine core comprising a compressor, a combustor and a turbine; atleast one propulsive fan for generating a mass flow of air to propel theaircraft; and an electrical energy store on board the aircraft, themethod comprising: supplying fuel to the engine core combustor toproduce a flow of combustion products for driving the engine coreturbine and compressor; monitoring conditions indicative of contrailformation; selectively altering the ratio of the mass flow of air by thepropulsive fan to the flow of combustion products of the combustor; andselectively assisting rotation of the engine core compressor by anelectric motor drawing energy from the electrical energy store duringthe altering of said ratio.

The altering of said ratio may comprise reducing the rate of fuelconsumption of the engine core and/or the rate of thrust/combustionproducts produced by the engine core. The rotation of the compressor maybe assisted when the rate of combustion and/or rotational speed of thecompressor reduce.

According to a third aspect of the invention, there is provided a datacarrier comprising machine-readable instructions for the operation of anaircraft propulsion controller to: receive sensor readings for aplurality of engine operation variables; monitor conditions pertainingto adverse contrail formation by combustion products from the engine;output control instructions in response to a determination of adversecontrail formation to alter a ratio of the mass flow of air by apropulsive fan to a flow of combustion products of an engine corecombustor; and concurrently output control instructions to selectivelyassist rotation of a compressor for the engine core by an electric motordrawing energy from an electrical energy store.

Any of the preferable features defined in relation to any one aspect ofthe invention may be applied to any further aspect. Accordingly theinvention may comprise various alternative configurations andcombinations of the features defined above.

BRIEF DESCRIPTION OF THE DRAWINGS

Practicable embodiments of the invention are described in further detailbelow by way of example only with reference to the accompanyingdrawings, of which:

FIG. 1 shows an overview of an aircraft having a distributed propulsionsystem;

FIG. 2 shows a longitudinal section through an aircraft engine accordingto one example of the invention;

FIG. 3 shows a longitudinal section through an aircraft engine accordingto a further example of the invention;

FIG. 4 is a graph depicting contrail factors for different operationalregimes of the engine of FIG. 2;

FIG. 5 is a simplified image of a gas turbine having a bypass with amodulating inlet;

FIG. 6 depicts components of a distributed propulsion system accordingto one example of the invention; and

FIG. 7 depicts components of a distributed propulsion system accordingto another example of the invention.

DETAILED DESCRIPTION OF THE DISCLOSURE

In FIG. 1 there is shown a schematic representation of an aircrafthaving an electrically assisted propulsive system 10 according to anexample of the present invention. Although the rest of the descriptionis mostly directed to aircraft having distributed propulsion units itwill be appreciated that distributed propulsion systems provide just onecontext in which the invention may be used and aircraft withoutdistributed propulsion may also operate in accordance with aspects ofthe invention described herein.

The electrical propulsive system 10 includes a plurality of electricalpropulsive units in the form of fans 12 which are rotatably driven byelectrical machines, for example superconducting electrical machines.

Each of the fans 12 includes a rotor having fan blades 14 mounted on arotatable hub and may have a blade pitch adjustment mechanism forsynchronously adjusting the pitch of the blades 14 relative to the airflow which passes them in use. Although only the propulsive units on thewings are shown as having blades 14, it will be appreciated that all ofthe propulsive units 12 include fans and blade arrangements. Theelectrical propulsive units 12 are placed in various locations aroundthe fuselage 22 and wings 20 of the aircraft, e.g. so as to ingestboundary layer air, which is energised and exhausted to providepropulsive thrust. Having a plurality of smaller propulsive units 12rather than two (or more) large gas turbine engines helps reduce dragand allows for a more efficient bypass ratio of the propulsive system,thereby increasing efficiency of the aircraft.

In the embodiment shown, there are six electrical propulsive units 12located towards the trailing edge of the wing and two located towards arear portion of the fuselage flanks. Two further units are located onthe tips of the wings. All of the propulsive units 12 may be of theducted variety or open rotor propellers as are known in the art.

The electrical machines which drive the propulsive units 12 may besuperconducting synchronous machines having superconducting rotors whichare permanently magnetised in use. The rotors are driven using statorwindings which may or may not be superconducting. Such machines areknown in the art. As will be appreciated, the electrical machines may beoperated as motors or generators. The term ‘electrical machine’ as usedherein is also intended to encompass a plurality of electrical machinesacting together to rotably drive (or be rotably driven in unison by) asingle rotor or rotor assembly, such as the propulsive fan of an engine.While a single electrical machine would typically be understood to drivea rotor/assembly near its hub or via a shaft, a plurality of electricalmachines could be arranged circumferentially around the rotor/assembly,and drive it from the circumference, for example via a ring.

The electrical propulsion system 10 also includes a plurality ofelectrical sources in the form of electrical generators 16, e.g.synchronous machines, which are driven by power plants, such as theshown two main gas turbine engines 18 located underneath the wings 20.The gas turbine engines 18 operate in a conventional manner but areconfigured for operation as part of a distributed propulsion system 10.In some examples, the engines 18 may have a reduced bypass ratio so asto reduce aero-dynamic drag, or the bypass duct is removed completely tofurther reduce the aero-dynamic drag.

The electrical generators 16 and the electrical machines of thepropulsive units 12 are electrically connected via cables in the form ofa bus system 26 and as such collectively form an electrical network. Theelectrical network may also include ancillary equipment in the form ofisolators and fault current limiters which are generally indicated byreference numeral 32. The electrical generators 16, bus system 26 andfault current limiters 32 within the network may or may not besuperconducting.

The propulsive system 10 includes an electrical energy store 30 and acontrol system which has at least one controller 28 which is configured,at least in part, to monitor and determine the required thrust for eachpropulsive unit. In some embodiments, the controller 28 may regulate thepitch of the fan blades to control the amount of propulsive forceproduced and to optimise the aerodynamic efficiency of the fan bladesover a wide range of combinations of aircraft forward speed and fanrotational velocity.

In the embodiment shown, an auxiliary power unit 24 is provided, e.g. atthe rear of the fuselage. The auxiliary power unit may for example be agas turbine, a reciprocating engine or a fuel cell. The auxiliary powerunit is arranged to provide electrical power to the aircraft when themain generators 16 are inoperable, for example, prior to the engines 18being started. The auxiliary power unit may be replaced or supplementedby an electrical storage device e.g. one or more battery units.

The separation of the power plant from the propulsive units means thatthe engine 18 is susceptible to contrail formation in a much wider rangeof atmospheric conditions since the exhaust contains all the watervapour released from combustion but does not contain all the heat energythat would be present in a turbofan exhaust since, in a distributedpropulsion arrangement, some or all of the fan work is exhaustedelsewhere. The invention may find particular application in such adistributed propulsion system but it is not limited thereto since anygas turbine engine may be susceptible to contrail formation undercertain operating and ambient conditions.

FIG. 2 depicts a simple schematic of an arrangement of a power plant 110for an aircraft, such as the engine 18 in FIG. 1, or else an engine fora non-distributed aircraft propulsion system. The power plant has anengine core that drives an electrical generator and a ducted bypass pastthe core, the core exhaust and the bypass exhaust mixing either in acombined duct or shortly downstream of the end of one or more of thecore or bypass duct. The engine core may be a reciprocating engine, or agas turbine. For the rest of this application the embodiments will bedescribed with respect to a gas turbine engine core.

The gas turbine engine comprises in axial flow series a series ofcompressors 106, a combustor 108 and a series of turbines 109. There isa general direction of airflow, indicated by arrow A, through theturbofan gas turbine engine in operation and the terms upstream anddownstream are used with reference to this general flow direction.

The flow through the engine flows through the compressors where it iscompressed and passed into the combustor 108 where it is mixed with fueland the fuel is burnt in the air in the combustor. The combustion of thefuel in the compressed air in the combustor 108 produces hot gasesincluding water vapour, which exit the combustor and flow downstreamthrough and thereby drive the turbines. The turbines drive thecompressors and an electrical generator 113 via shafts 111, whichdrivingly connect the turbines with the compressors and the electricalgenerator. The electrical generator 113 is driven in this exampleselectively by either or both the engine core (high pressure) turbineand the low pressure turbines 109 via respective shafts. In otherexamples a separate generator could be provided for each shaft as willbe described in further detail below.

The exhaust gases leaving the turbines flow through the exhaust nozzleassembly to provide some propulsive thrust. At, or just after theexhaust nozzle, air passing through the bypass 107 is mixed with thecore exhaust. The air that passes through the bypass duct has, at theexit of the duct, a higher temperature than the ambient air. Thecombined mixed exhaust flow has a significantly lower contrail factorthan the engine core exhaust flow alone. By increasing the rate of heatinput to the flow of air and/or the mass flow rate of air through thebypass duct relative to the rate of water vapour input into or formedwithin in the core, the contrail factor of the combined mixed exhaustcan be further reduced.

A two-shaft, or two-spool, engine arrangement is shown in FIG. 2 inwhich a low pressure compressor (i.e. comprising a plurality ofcompressor rows/stages) feeding air to the engine core is provided onthe lower pressure shaft between the fan 102 and the turbine(s) 109.Thus the low pressure compressor acts as a booster in a so-calledboosted turbofan configuration.

FIG. 3 shows another engine example, for which corresponding numeralshave been used to denote corresponding features with FIG. 2. The exampleof FIG. 3 differs to that of FIG. 2 in that a gearing arrangement 118(e.g. a reduction gearbox) is provided in the force path between thelower pressure shaft of the shafts 111 and the fan 102. This is to allowthe fan 102 to rotate at a more optimal speed relative to that of thecorresponding turbine and/or intermediate pressure compressor (IPC). Itwill be appreciated that, in such a geared turbofan configuration, thecompressor 106 associated with the radially inner shaft 111 is typicallyreferred to as the IPC, rather than the low pressure compressor of FIG.2. The gearbox 118 may be provided between the IPC and the fan 102 suchthat the fan is driven via the gearbox by the corresponding intermediatepressure shaft. FIG. 4 is an exemplary graph showing the mixing line ofa fan exhaust 40, a core exhaust 42 and an exhaust 44 comprising thecombination of the core exhaust and the fan exhaust. Each of the mixinglines 40, 42, 44 is characterised by a gradient, sometimes known as the“contrail factor”, which is the ratio between the added water vapour toadded heat in the exhaust relative to ambient conditions. The lower leftend of each line 40, 42, 44 represents ambient conditions. Each mixingline shows the evolution of temperature and water vapour partialpressure from exhaust conditions to ambient conditions. Also shown isthe water saturation curve 46 of air at a given temperature. If any partof a mixing line lies on or to the left of the water saturation curvethere is a likelihood that contrails will form or may form. If a mixingline lies entirely to the right of the water saturation curve there isno chance, or minimal chance, that contrails will form.

Turning first to the core exhaust, the gradient of the line 42 isrelatively steep i.e. although the exhaust is hot there is also a largeamount of water vapour within the exhaust. Within some regions ofoperation at least part of the line 42 lies to the left of the watersaturation curve indicating that there is a risk of contrail formation.The contrail factor of the fan (being the gradient of the line 40), bycontrast, is close to zero. Although the fan inputs work into the flowthis is achieved without any water being supplied into the flow. At nopoint does the line pass to the left of the water saturation curve sothere is no, or minimal risk of contrail formation. The contrail factor,and hence the gradient of the mixing line 44, of the mixed exhaust isless than the contrail factor, and hence the gradient of the mixing line42, of the core exhaust but greater than the contrail factor, and hencethe gradient of the mixing line 40, of the fan exhaust. It will beappreciated that, by increasing the temperature of the fan flow, orincreasing the air mass flow rate through the bypass duct 107 relativeto the air mass flow rate through the engine core and, assuming that therate of water mass flow within the mixed exhaust stays the same, ordecreases, the contrail factor can be further reduced to the point atwhich the risk of contrail formation is negated for all but extremesituations.

A decrease in the rate of water vapour emission can be achieved byreducing the rate of fuel-flow to the combustor 108. Thus, vapouremission and/or the mixing line may be used for determining when and towhat extent to control the rate of fuel-flow to the combustor 108 forcontrail suppression.

As mentioned earlier, in a distributed propulsion architecture thecoaxial fan 102 may not be needed to provide instantaneous propulsivethrust as this is achieved by the distributed fans. Accordingly, thecoaxial fan 102 may be used selectively and it may be desirable toselectively open and close the bypass in accordance with the need foroperation of the coaxial fan 102.

At times when contrail suppression, or fan operation, is not requiredthe airflow through the bypass may be reduced, or optionally closed off,in order to reduce drag as far as possible. FIG. 5 depicts a simplifiedimage of a possible arrangement with FIG. 5a depicting an operationalstate in which the bypass is closed and FIG. 5b depicting an operationalstate in which the bypass is open.

The bypass closure 119 may be one or more flaps that may be movedbetween a position in which the bypass is closed (FIG. 5a ) orsubstantially closed and a position in which the bypass is open (FIG. 5b). In the position in which the bypass is closed the flap preferablydirects the air radially outwardly and around the engine nacelle 117 bypresenting a sloped surface to the air flow. In an alternativearrangement, the flap provides a bluff surface but this will have adetrimental impact on the drag of the engine.

The flap may be hinged at a forward edge in which it lies flush againstthe radially inner wall 115 of the bypass when the duct is open. Theflap may be hinged at a rearward edge in which the flap rotatesoutwardly to form an extension of the nacelle or radially outer wall 117of the bypass duct when the duct is open.

Alternatively, the bypass may be closed by an alternative mechanism e.g.an inflatable bag that may lie against a surface of the bypass duct inan uninflated state when the bypass duct is required to be open andacross the bypass in an inflated state when the duct is required to beclosed. The bag may be shaped such that in an inflated configuration itpresents a slope to the air flow that directs air radially outwardly andaround the engine.

In examples of variable bypass closures shown in FIGS. 2 and 3, theexhaust nozzle assembly 116 comprises two concentric exhaust nozzles, aradially outer bypass, or fan, exhaust nozzle 112 and a radially innercore exhaust nozzle 114. The core exhaust nozzle 114 is defined at itsradially outer extent by a generally frusto-conical core nozzle wall 115and at its radially inner extent by a frusto-conical engine plugstructure 122.

The bypass, or fan, exhaust nozzle 112 is defined at its radially outerextent by a nacelle, or fan casing, 117 and at its radially inner extentby the core nozzle wall 115. The bypass, or fan 102, exhaust nozzle 112may be a variable area fan exhaust nozzle.

The bypass duct 107 is also defined at its radially outer extent by thenacelle, or casing, 117, which is generally annular and arranged coaxialwith the engine axis 1. Thus, the nacelle, or fan casing, 117 defines aflow passage through the turbofan gas turbine engine 110. The bypass, orfan, exhaust nozzle 112 is arranged at the downstream end of the nacelle117.

A controller 120 may be arranged to control an actuator 132 to vary thecross-sectional area of the variable area fan exhaust nozzle 112according to sensor signals. The nozzle may adopt different positions160A, 160B as required. Although in the above-described example, theactuator 132 has discrete positions of actuation, it will be appreciatedthat any flow opening which is actuable according to the invention willtypically be variably actuable over the available range of actuation soas to adopt any suitable condition within that range according to thesensed operating/ambient conditions. Accordingly the actuator or flowopening may be range-taking. In one example a plurality of predeterminedactuator positions and/or flow opening areas may be defined such thatthe controller selects one of the predetermined options in use, forexample the predetermined option which is closest to an optimal positiondetermined by the controller.

Whilst the above description and FIGS. 2, 3 and 5 describe the variablecontrol of the bypass flow for completeness, it will be appreciated thatthose details concern certain implementations of the invention only, andthe associated features may be excluded from other examples of theinvention in which they are not required.

In an alternative embodiment a fan may be associated with the bypassduct and may be operable or operated at a different rotational velocityat times where the contrail suppression is deemed necessary anddesirable. The rotational velocity of the fan may be controlled tochange either or both of the heat input into the flow of air through thebypass duct or the mass flow rate of air passing through the bypass. Thechange in rotational velocity may feature in combination with otherfeatures such as a variable pitch rotor, variable area nozzle and/orvariable outlet guide vanes 119 to help accommodate a wide range of fanpressure ratios and aircraft forward speed.

The rotational velocity of the fan may be used to capture energy duringdescent, acting as a turbine and driving the generator 113 whichreplenishes an energy store and/or provides energy to aircraftelectrical systems.

The fan may be driven electrically, mechanically or as a hybrid wherethe fan is driven mechanically with a supplemental electrical drive, orelectrically with a supplemental mechanical drive. A mechanical drive(e.g. a co-axial fan 102 with the engine core as shown in FIG. 2) with asupplemental electrical drive may be used where the fan is used forgeneral propulsive purposes as well as for contrail suppression.

When the engine is operating in a mode that suppresses contrails thebypass air flow may be increased by increasing the fan rotational speedwith a possible change in fan blade pitch and/or a possible change infan OGV pitch and/or a change in bypass nozzle area.

In an arrangement which supplements a mechanically driven fan, or wherethe fan is purely electrically driven, the core fuel flow, core massflow, and hence the water-vapour emissions in the core, can be decreasedand an electrical drive to the fan used to supplement the reducedmechanical drive.

As there is a reduction in core mass-flow and water input into the coremass-flow there is an improved contrail factor even if the electricaldrive is used to input enough power to the fan to maintain the desiredoperating thrust. By increasing the electrical drive for the fan to pusheven more air through the bypass duct than required for the operatingpropulsion the dilution effect can be enhanced and hence the contrailfactor can be further reduced. The thrust produced by the otherpropulsive fans (if present) may be adjusted by the system controller 28to maintain the commanded or desired total thrust required by theaircraft. Conversely, in circumstances where the available power of themotor 113 is insufficient to maintain the desired level of thrust fromthe engine's co-located fan 102 during a contrail-suppression conditionin which the core fuel-flow rate has been reduced, the thrust from theother propulsive fans 12 may temporarily be increased to compensate fora reduced level of thrust from the co-located fan 102 during thecontrail-suppression condition.

FIG. 6 depicts a simplified image of the propulsion system that has anenergy storage unit 200 which may be a battery or other form of energystorage deemed suitable such as, for example, flywheel, capacitor-array,compressed air, liquid air energy store. An engine 210 comprises acoaxial fan 212, an engine core 216 and an electrical machine (i.e. amotor/generator) 214 configured selectively to drive or be driven by thefan 212 and/or the engine core 216. The electrical machine 214 isdrivingly connected to the fan and/or engine core via one or moreshafts.

The engine 210 and its associated engine core 216 may be either of theengines 110 shown in FIGS. 2 and 3. The electrical machine 214 may thusbe the motor/generator 113 shown in those figures.

The engine core 216 mechanically drives the fan 212 and/or electricalmachine 214 when used as a generator.

A remote propulsive fan 220 comprises a fan 222 and an electricalmachine (i.e. a motor/generator) 224 that is either integral with ordrivingly connected, e.g. via a shaft, to the fan and configured todrive or be driven by the fan according to a prevailing mode ofoperation. In practice, there may be more than one remote propulsivefan. In some embodiments, there could be many small remote propulsivefans (each with its own motor/generator) configured to ingest andre-energise the boundary layer air flowing over or under the aircraft'swings or other surfaces such as the fuselage.

A first electrical cable 230 is configured to transport electricalenergy (in either direction, according to the prevailing mode ofoperation) between the energy storage unit 200 and the motor/generator224. A second electrical cable 232 is configured to transport electricalenergy (in either direction, according the prevailing mode of operation)between themotor/generator 214 and the energy storage unit 200.

A third electrical cable 234 is configured to transport electricalenergy between the motor/generator 214 and the motor/generator 224. Thedirection of energy flow would, in all envisaged modes of operation, befrom 214 to 224. However, flow in the opposite direction should not beruled out.

There would also be a control and decision/making unit (not shown) whichwould be in signal communication with the energy storage unit 200, theengine core 216, the motor/generator 214 and the motor/generator 224.The control and decision making unit could be an integral part of, or amodule within, an engine-control-unit (ECU) of the engine 210, to bedescribed below, or it may be a separate physical entity in signalcommunication with the ECU. If required, the control and decision makingunit may also be in signal communication with the co-located fan 212,and the fan 222 (if present), for the purposes of instructing changes totheir respective blade pitches. Furthermore, thecontrol-and-decision-making unit could also be in signal communicationwith any variable-area nozzles which may be present, in order todetermine and instruct changes thereto.

In FIG. 6, a gearbox 215 unit is provided to change the rotational speedratio of the fan and/or the HP shaft/compressor of the engine corerelative to the rotor of the electrical machine 214. The gearbox unit215 may comprise a selector or clutch arrangement for selective couplingto either or both of the fan rotor/shaft and engine core compressor.When used with an electric motor to drive both the fan 212 and enginecore 216 shafts, the gearbox may have one or more predetermined gearratio for driving both shafts at fixed relative rotational speeds.Typically a plurality of gear ratios will be provided such that asuitable gear ratio can be selected by the control unit that mostclosely matches the desired relative rotational speeds for the enginecore compressor and the booster or IPC rotor, depending on the specificengine configuration.

In FIG. 7, the single electrical machine configuration for the engine210 has been replaced with a dual-electrical-machine configuration. Thusa first electrical machine 214 is provided for the engine core 216 and asecond electrical machine 217 is provided for the fan 212. In thisconfiguration, each electrical machine 214, 217 may have its ownrespective gearbox. This may reduce the complexity of any gearboxrequired and may allow each electrical machine to be controlled with agreater degree of freedom from the operational requirements of theother. Thus the speed of each electrical machine when operating aseither a generator or motor can be tailored to an optimal speed for theshaft to which it is connected under the given operational conditionsinstead of being limited to a number of fixed relative gear ratios. Forexample, when providing electrical assistance to both the fan 212 andengine core 216, it may be desirable to have greater freedom over therotational speed of the fan 212 compared to that of the engine core 216.

In the configuration of FIG. 7, the motor generator 113 of FIGS. 2 and 3would be coupled to only one of shafts 111 and would thus correspond toone of electrical machines 214 and 217, with the other electricalmachine being coupled to the other shaft 111.

For different operational states/requirements of the engine, ordifferent flight regimes, the propulsion system may be controlled in adifferent manner. For example, when the thrust demand for the propulsionsystem is relatively high, the engine 210 will typically be running withhigh power output and powering its own coaxial fan 212. If additionalthrust is required, i.e. beyond the output of the engine core, theenergy storage unit could supplement the power to the fan by driving theelectrical machine(s) 214/217 as motors. Additionally or alternatively,the energy storage unit could provide supplementary power to the remotepropulsive fan 220 (if present). If additional thrust is not required,the engine core could itself drive either or both of the electricalmachines 214/217 as generators.

Such a high thrust demand could apply during take-off and climb(including in particular step-climb and top-of-climb as well as theclimb away from a departure airport), or during certain aircraftmanoeuvres, e.g. such as emergency manoeuvres.

In contrast, when the thrust demand is very low, engine 210 may be runat an idle speed or may be shut down for a period of time, e.g. for acruise phase of a short flight. Thrust could be provided predominantlyor entirely using electrical energy from the energy storage unit 200 todrive the co-located fan 212 within the engine 210, and optionally theremote propulsive fan 220.

During cruise, the energy storage unit could also be receivingelectrical energy from the generator 214/217 of the engine 210, and thusbe charged. The engine throttle is set at a suitable level to meet theaircraft's thrust requirement as well as providing charge to the energystorage unit 200.

In the descent phase of flight, the engine 210 is in its unlit or idlestate. The fan 222 of the propulsive fan 220 (if present) is beingdriven (“windmilling”) by air-flow resulting from the forward motion ofthe aircraft, and as a result is causing the generator 224 to generateelectrical energy which is used to charge the energy storage unit 200.Additionally or alternatively, the co-located fan 212 of the engine 210can be used in a “wind-milling” capacity to charge the energy storageunit 200 with electrical energy generated by the generator 214/217. Thiswould advantageously allow aircraft speed to be regulated in support ofsteeper descents without the need to configure the aerofoils of theaircraft wing into a high-drag high-noise configuration. Alternativelythe airflow to the co-located fan 212 could be closed off (as shown inFIG. 5) to reduce drag.

Regardless of the normal operation modes of the propulsion system, acontrail-suppression mode is also available. This mode is selected whena) contrail formation is observed and/or predicted according to observedambient conditions and the engine's operating condition prior toselection of this mode, and optionally b) ambient conditions areconducive to contrail persistence. Further criteria could be employed indeciding whether or not to suppress the formation of a contrail.

Whilst the term ‘contrail suppression’ is used herein, it is intendedthat partial or complete suppression of a contrail is contemplated bythe invention. If complete suppression of contrail formation is notachieved during an instance of use, the invention may nonetheless allowa reduction in the optical depth of the formed contrail, e.g. due toactivation of a smaller proportion of the potential condensation nucleiin the engine exhaust. Even where complete suppression of a contrail isachievable, it may be undesirable due to other control considerationsand so control settings that enable partial contrail suppression may beenacted.

During contrail suppression the fuel-flow rate to the engine core 216 isreduced and may for example be reduced down to zero (or very close tozero corresponding to an idle setting). In this mode of operation, theaircraft may fly exclusively on stored electrical energy drawn from theenergy storage unit 200, for the duration of the contrail suppressionrequirement. Thus the contrail suppression mode may resemble anall-electric cruise mode as described above. This mode is selected if amore modest reduction in engine fuel flow rate is insufficient tosuppress contrail formation, or if electric-only cruise is deemedadvantageous for other (i.e. non-contrail-related) reasons.

When the fuel flow rate is non-zero, i.e. when the engine is operationalbut generating significantly less thrust than would be required topropel the aircraft without electrical assistance, the potential rangeof relative rotation speeds between the engine core shafts 111 (in FIG.2) is large. That is to say, the fan 102/212 could be driven at varyingrotational speeds to provide thrust, whilst the engine core compressorand turbine are run at lower than normal speeds.

This ability to significantly alter the relative rotation speeds betweenthe engine shafts allows greater freedom of control but places theengine core in a flow regime that is outside the scope of aconventional, i.e. electrically-unassisted, engine. The available rangeof control could thus force the booster/IPC into one or more unwantedoperational regimes, such as an engine/compressor stall or surgecondition. This is because the reduction in engine core speed causes areduction in ingested mass flow at the engine core compressor and thus areduction in mass flow for the booster/IPC. However because thebooster/IPC is associated with the fan it can be driven at variablespeed according to a desired thrust level to be supplied by the fan,rather than its own requirements.

It has been determined that a full range of propulsion system operation,e.g. for reasons of contrail suppression and/or engine efficiency, maybe permitted by selectively assisting the engine core 216 via the motor214 when required to maintain safe engine operation. For example, inperiods when the engine core speed is reduced and the fan 212 is drivenwhether solely or partly by its electric motor, the engine corecompressor could also be selectively driven by the electric energy store200 via motor 214 when the IPC or booster is facing an elevated risk ofstall/surge.

Returning to FIG. 2, the turbofan gas turbine engine 110 has a pluralityof engine operation sensors shown schematically at 124 and 126. Thesensor examples 124 and 126 may be arranged to measure pressure at theintake 105 (i.e. upstream of the fan 102) and also the total pressure inthe bypass duct 107, thereby allowing determination of the powerproduced by the engine 110. However this provides just one example of anengine operation sensor arrangement and there are many additional oralternative engine sensor arrangements that may be used in conjunctionwith the invention as will be described below.

A controller 120 is also arranged to receive signals of sensedparameters from externally of the engine 110, such as from one or moreambient condition sensor 128 and/or a contrail detection sensor 130. Theambient sensor 128 comprises a plurality of sensors for measuringaltitude (e.g. ambient pressure), temperature and/or humidity. Thecontrail detection sensor in this example comprises an optical depthsensor having a field of view downstream of the engine exhaust (i.e. todetect formation of contrails aft of the engine). Alternative contraildetection sensors could be used, such as an acoustictransmitter/receiver. Any of the external sensors 128, 130 may bemounted on the aircraft body or wing.

The controller 120 is arranged to receive signals of sensed parametersin use from the engine operation and other sensors. The sensors 124-130supply signals/measurements to the controller 120 via connecting leads(e.g. electrical or fibre-optic cable) or else a local wireless network.

The controller 120 is typically a general engine or propulsion systemcontroller, i.e. arranged to control numerous aspects of engine/systemoperation in addition to the contrail mitigation system describedherein. Accordingly, it will be appreciated that such a controller willtypically receive inputs from a significant number of other sensors.However, one or more bespoke control algorithm (e.g. one or more moduleof code) will control operation of the controller as a contrailmitigation controller in accordance with the invention. Suchfunctionality dictates which engine operation variables are availablefor control in accordance with the contrail mitigation strategy. Thecontroller will also determine what scope of control is available forthe purpose of contrail mitigation based upon other engine operationrequirements such as engine-throttle-setting and/or thrust-requirementsetting that must be prioritized above contrail mitigation needs.

Ambient condition sensors 128 are as described above and are arranged tooutput real-time or near-real-time information concerning the ambientpressure, ambient temperature and/or ambient humidity of air throughwhich the aircraft is flying. In another example, one or more of thosesensors 128 may optionally be replaced and/or supplemented by a databaseof forecast data (e.g. obtained prior to the flight and/or updatedperiodically during the flight via radio-link or similar) which detailsthe ambient condition(s) that the flight will encounter along itsproposed flight routing and altitude profile, taking account of theexpected time at which each point of the proposed route will be passed.

A deployment policy or contrail suppression policy may be employed. Thepolicy allows prioritisation of the control system to achieve greatestbeneficial climate impact. That is to say the policy determines how bestthe invention should be employed to ensure that, within the confines ofsafe engine operation, any increased fuel burn due to reduction ofoverall propulsion system efficiency is outweighed by the beneficialimpact of contrail suppression/mitigation. The policy can be implementedby one or more routine or calculation to assess whether or not toattempt contrail suppression according to either or both of:

-   -   An assessment of the likely climate warming impact if the        contrail is allowed to form    -   An assessment of any cost penalty associated with allowing the        contrail to form

The assessment of climate warming impact would primarily take intoaccount one or more of the following factors:

-   -   Will the contrail persist?    -   Over the contrail's lifetime/persistence, what is the balance        between day-time and night-time?    -   What is the temperature of the ambient air in which the contrail        will reside?

Using answers to the above questions, a decision on whether or not todeploy the contrail mitigation/suppression mode can be taken, e.g.according to one or more contrail characteristic, such as whether or nota predetermined threshold of climate-warming impact of the contrail ismet or exceeded.

In a decision-making process for determining whether control steps arerequired to mitigate against contrail formation or else promoteoperating efficiency, the control system operates a primary loop tocheck whether the operating conditions (i.e. the ambient conditionsand/or engine operation parameters) have changed materially since aprevious iteration. The operating conditions could comprise any, or anycombination of, ambient temperature, ambient pressure, ambient humidity,altitude, and/or engine throttle setting or engine operating point. Anymaterial change may comprise change of any one or more parameter to anextent that will materially alter a contrail characteristic (orpersistence, or occurrence of contrails). This primary loop can beiterated without changing existing settings until a relevant change tooperating conditions is determined. Current operating conditions couldbe stored at each iteration such that the next iteration can comparecurrent operating conditions against one or more recorded set ofprevious operating conditions. Alternatively, current operatingconditions are stored during the first iteration of the primary loop andthereafter only when a material change to the operating conditions isdetected relative to the previously stored value. In any example, a logof previous operating conditions is maintained to allow comparison withcurrent or most recently sensed conditions.

The controller then determines, e.g. according to a currentusage/deployment policy, whether it is deemed appropriate to use amethod for mitigating contrail formation under the current operatingconditions. For instance, it may or may not be considered appropriate toonly use this invention when contrails persist, e.g. if ambient relativehumidity over ice is 100% or greater.

When operating at reduced engine core speeds and/or modified fan speeds,the controller determines whether the intended engine core operatingconditions induce an adverse operating state for the engine. The surgemargin may be taken into consideration when determining an adverseoperating state, although the invention need not be limited only tosurge risk monitoring. Various options of individual thresholds forindicating adverse operating states for the engine core could beselected, such as pressure change across the compressor, flow rate,shaft rotation speed, torque, combustion temperature, fuel flow rate orany combination of those parameters.

It will be appreciated that the dynamic nature of the system willrequire thresholds for deciding when selective electrical motorassistance to the engine core compressor is required to be set based onrelative values for a plurality of parameters. For example a combinationof pressure drop and flow rate may be used. Either or both of thoseparameters may be used in conjunction with the torque applied to eitheror both engine shafts 111.

The relevant threshold(s) will be selected with safety margins appliedbased on known operational characteristics of the relevant compressor,e.g. the compressor map.

The safety margin may also account for any current rate of decelerationof the compressor and/or the normal time required to enact a desiredspeed change for the compressor and the associated mass flow ratethrough the compressor. Thus the controller may operate according to apredictive algorithm to avoid onset of potential surge conditions ratherthan actual measured surge onset.

If an intended adverse operating condition is determined, the controllerdetermines the degree of electrical assistance to the engine corecompressor that is required to avoid the onset of the adverse condition.The controller also determines whether sufficient electrical energy isavailable to avoid onset of the adverse condition, either in a transientmanner or else in an ongoing/prolonged manner. In this way, prior toinstructing a change in engine operation, the controller can determinewhether the change can be made without compromising a safety conditionor margin for the required duration.

If the relevant adverse condition mitigation requirements can be met,the controller can instruct the change in operating conditions, alongwith any electrical assistance to the engine core (i.e. high pressure)spool as required.

An assessment of whether or not a particular contrail characteristicwould warrant mitigation steps is undertaken by checking whether one ormore ambient sensor reading and/or engine operation parameter achieve athreshold level. In one example, a vapour trail detection sensor 130 maybe used to trigger contrail mitigation action and/or to verify adetermination of the presence or absence of a contrail. If there hasbeen no material change in the operating conditions, or a materialchange in the operating conditions is detected but any of the otherconditions described above are not satisfied, then no change to thecurrent contrail mitigation control settings are made.

In various examples of the present invention, it is considered pertinentto optionally disable the contrail suppression system during one or moreflight phase, such as during take-off, climb and/or approach, whereemergency situations may demand that more thrust is commanded quickly.The control system may operate according to different operationalpriorities during these conditions so that the electrical devices areready to provide additional power promptly if and when it is requestedby the pilot (e.g. in case of emergencies etc.).

The present invention may focus particularly on the avoidance ofcontrail formation in regions of ice-super-saturated (ISS) air. However,on average, aircraft spend only a small proportion of their flight-timein ISS air, and so the proportion of the flight during which a contrailsuppression mechanism needs to be active in order to mitigate against amajority of the negative climate impact of contrail formation isrelatively small. Thus, instead of attempting to suppress all contrailsto some level, the invention focuses on the careful identification ofspecific periods only in which suppression of contrail formation oradvantageous modification of contrail properties can bring about a netpositive climate impact beyond what would be achieved using conventionalengine control.

Whilst the embodiment described above refers primarily to an arrangementin which an engine comprising a core and a fan forms part of adistributed propulsion system comprising remote propulsive fans, it willbe appreciated that the contrail-suppression method comprising areduction in the core fuel-flow rate of an engine in conjunction withelectrical-supplementation of the power supplied to the fan of the sameengine could also be applied to a gas turbine, including ahigh-bypass-ratio gas turbine, irrespective of the existence of remotepropulsive fans.

It will also be appreciated that a related approach to contrailsuppression, effective particularly at low throttle conditions or idleconditions such as during descent or during loitering, involves atemporary reduction in electrical power offtake from the engine, meetingsome or all of the aircraft's electrical power requirements (such asavionics, aileron and elevator control, air-conditioning etc.) fromstored energy rather than from the engine power offtake. This is becausethe engine power offtake slightly reduces the exhaust temperature, butwithout altering the water vapour emission rate. A reduction in powerofftake from the engine thus reduces the engine's contrail factor, inother words the gradient of the mixing line, and hence reducessusceptibility to contrail formation. The materiality of this effect isgreater when the magnitude of the power offtake is a significantfraction of the rate of fuel-energy released during combustion, in otherwords at low throttle settings. Optionally, mixing between the coreexhaust and the bypass exhaust of the engine could be enhanced by alobed mixer, which might have the additional benefit of slightlyincreasing propulsive efficiency.

Whilst the invention is described particularly in the context ofcontrail suppression, it will be appreciated that the invention mayequally apply to propulsion system in which fuel/energy efficiency isprioritised over, or instead of, contrail formation. Accordingly, therange of control permitted by use of the electrical machines selectivelyto drive either or both of the high pressure and low pressurecompressors as described herein may be useful even when contrailsuppression is not required, for example to avoid adverse operatingstates, for example during engine deceleration.

The invention claimed is:
 1. An aircraft propulsion system comprising:an engine having a propulsive fan and an engine core, the engine corecomprising a compressor, a combustor and a turbine driven by a flow ofcombustion products of the combustor, the propulsive fan configured togenerate a mass flow of air that bypasses the engine core and propels anaircraft on which the engine is located; an electrical energy store onboard the aircraft; at least one electric motor arranged to drive thepropulsive fan and the engine core compressor; and a controller arrangedfor control of the at least one electric motor to mitigate creation of acontrail caused by the combustion products by altering a ratio of themass flow of air by the propulsive fan to the flow of combustionproducts of the combustor, wherein control of the at least one electricmotor comprises selectively and concurrently driving both the propulsivefan and engine core compressor.
 2. The aircraft propulsion systemaccording to claim 1, wherein the at least one electric motor isconfigured to selectively assist the engine core compressor bysupplementing torque applied to the compressor via the turbine due tothe engine core combustion process.
 3. The aircraft propulsion systemaccording to claim 1, wherein the controller is configured to monitorconditions indicative of contrail formation and the controller isconfigured to reduce the flow of combustion products produced by theengine core to mitigate or suppress a creation of a contrail and thecontroller is further configured to concurrently increase power suppliedto the propulsive fan by the electric motor in order to meet a thrustdemand for the aircraft.
 4. The aircraft propulsion system according toclaim 1, wherein the engine comprises a further compressor and thecontroller controls operation of the at least one electric motor todrive the engine core compressor at a rotational speed to achieve atleast a threshold value of an operational variable for the furthercompressor.
 5. The aircraft propulsion system according to claim 4,wherein the operational variable comprises an absolute or relativeoperational variable between the engine core compressor and thepropulsive fan or further compressor.
 6. The aircraft propulsion systemaccording to claim 4, wherein the controller controls operation of theat least one electric motor so as to prevent operation of the furthercompressor from crossing a predetermined threshold value for acompressor surge margin.
 7. The aircraft propulsion system according toclaim 4, wherein the further compressor comprises a booster orintermediate pressure compressor arranged to be driven with thepropulsive fan by the at least one electric motor.
 8. The aircraftpropulsion system according to claim 1, wherein the at least oneelectric motor is controlled according to a control hierarchy whereinpriority is given to driving the propulsive fan for contrail mitigationand the engine core compressor is selectively driven by the electricmotor so as to maintain a safe engine operating profile.
 9. The aircraftpropulsion system according to claim 1, wherein the at least oneelectric motor comprises a first electric motor arranged to drive thepropulsive fan and a second electric motor arranged to drive the enginecore compressor.
 10. The aircraft propulsion system according to claim1, wherein a first electric motor drives both the propulsive fan and theengine core compressor via a selective transmission system.
 11. Theaircraft propulsion system according to claim 1, wherein the one or moreelectric motor drives the engine core compressor when the propulsive fanis concurrently driven by the at least one electric motor.
 12. Theaircraft propulsion system according to claim 1, wherein the propulsivefan comprises a booster.
 13. The aircraft propulsion system according toclaim 1, wherein the propulsive fan is driven by an intermediatepressure shaft of the engine, and a gearing is provided for altering arotational speed of the propulsive fan relative to a speed of theintermediate pressure shaft.
 14. The aircraft propulsion systemaccording to claim 1, wherein a gearbox is provided in a force pathbetween the at least one electric motor and the propulsive fan and/orthe compressor of the engine core.
 15. A method of operating an aircraftpropulsion system having an engine including a propulsive fan and anengine core, the engine core comprising a compressor, a combustor and aturbine, the propulsive fan configured to generate a mass flow of airthat bypasses the engine core and propels an aircraft on which theengine is located; and an electrical energy store on board the aircraft,the method comprising: supplying fuel to the combustor to produce a flowof combustion products for driving the turbine and the compressor;monitoring conditions indicative of contrail formation; selectivelyaltering a ratio of the mass flow of air by the propulsive fan to theflow of combustion products of the combustor; and concurrently assistingrotation of the compressor by the electrical energy store during thealtering of said ratio.
 16. A data carrier comprising machine-readableinstructions for operation of an aircraft propulsion controller to:receive sensor readings for a plurality of engine operation variables;monitor conditions pertaining to adverse contrail formation bycombustion products from an engine core combustor; output controlinstructions in response to a determination of adverse contrailformation to alter a ratio of a mass flow of air by a propulsive fan toa flow of combustion products of the engine core combustor, thepropulsive fan included in an engine comprising the engine corecombustor, the propulsive fan configured to generate the mass flow ofair that bypasses an engine core and propels an aircraft on which theengine is located, the engine core including the engine core combustor;and concurrently output control instructions to selectively assistrotation of a compressor for the engine core by an electrical energystore.
 17. An aircraft propulsion system comprising: an engine having apropulsive fan and an engine core, the engine core comprising acompressor, a combustor and a turbine driven by a flow of combustionproducts of the combustor, the propulsive fan configured to generate amass flow of air, which bypasses the engine core and propels an aircrafton which the engine is located; an electrical energy store on board theaircraft; at least one electric motor arranged to drive the propulsivefan and the engine core compressor; and a controller arranged forcontrol of the at least one electric motor to alter a ratio of the massflow of air by the propulsive fan to the flow of combustion products ofthe combustor, wherein control of the at least one electric motorcomprises selectively and concurrently driving both the propulsive fanand engine core compressor.
 18. The aircraft propulsion system accordingto claim 5, wherein the operational variable comprises rotational speed,torque, flow rate, and/or pressure drop.